Since the Direct baseline architecture is 2 J-246's, with two JUS's, one fully fueled, one fueled just enough for LEO of CEV/LSAM.

Could both JUS's be partially fueled, and have a rendevous in LE1 or LLO? You have to expend two anyway, couldn't the docking procedures be simplified that way?

Also, in that vein, if a Centaur or D4US (are they the same? Don't know much about them, if they are two terms for the same upper stage or not) can take Orion around the moon for an Apollo 8 type flyby, could one be used to take the LSAM there too, again for a LE1 or LLO rendevous?I know fueled LSAM is heavier than Orion, so I don't know.But now I'm curious.

You've just stumbled upon what, in my view, is the biggest weakness of the EOR-LOR mission profile. Try as you might, you can't distribute the total payload evenly among multiple identical launch vehicles without on-orbit propellant transfer.

If the DIRECT ethos can be distilled to "one kind of vehicle launched multiple times", then EOR-LOR is a questionable mission profile, a holdover from the 1.5-launch approach that doesn't make as much sense in a 2-launch architecture.

The DIRECT 3.0 architecture calls for one launch of about 100 mT and another of about 70 mT. It's not really a 2-launch architecture, hence the heavily-offloaded J-24x CLV and the barely-viable J-130 CLV alternative.

The J-246 EDS baseball card gives a 79mT payload through TLI.

So a 79mT CLV lift can be accomodated.

The J-246 CLV baseball card can only lift an 84mT payload to orbit with appropriate levels of safety for a crew (ie "with additional 10% Reserve"). The notional payload for the crewed vehicle is a lot higher than it appears to be because the margin is effectively also payload that needs to be accounted for.

5mT unused CLV lift is as close as you'll ever get to two fully-utilised identical vehicles.

Which reminds me -- I just heard that NASA has gone and produced a new "Analysis" into DIRECT 3.0 in which the same old BS is apparently doing the rounds -- yet again.

While I've only heard a few details, they're b*tching about the pmf being something which NASA is unable to build (yeah, we know that, which is precisely why we want Industry to handle that element because of their much greater experience instead, duh!).

And they're trying to depict the architecture as only being able to send a 29mT lander through TLI.{snip}

It sounds like the Direct team needs to follow the political parties and set up a Rapid Rebuttal Group.

I am not looking here at developing 2 Jupiter cores. What I am suggesting is to develop a lighther core which could be used with 3 SSME engines and an engine plug for the J-130. The same core could be used with 4 SSME engines, two vertical support beams and an upper stage to create a J-24x. The idea here is to permit the use of the J-130 for the lunar CEV+LSAM flight. The EDS flight on a J-24x is not the problem here as it has plenty of margin compared to the J-130 CLV flight.

It seems likely to me that the core + two beams stage will be less mass efficient than a core stage optimized for carrying an upper stage plus payload. This is not the way to go. Direct has it right. You need the highest performance on the J-24x vehicle, so you optimize for that. So what if the J-130 isn't optimized? It's not the vehicle that needs the highest performance.

That's a pretty concise summary of the situation. As I've argued before, RS-68 and J-2X are disappointing engines whose development has come at the cost of fielding much more promising engines such as RS-84 and RL-60. Delta IV will probably be the first and last vehicle to use RS-68, and J-2X (aka Vulcain reinvented for the NIH-afflicted) will probably never fly.

Eeek too many different types of engines to keep track of! I'm wishing I had a scatter plot with thrust on the horizontal axis (log scale), ISP on the vertical axis and one labeled point per engine, with symbol type denoting propellants. Does such a plot already exist or should I make it myself?

Edit: I just made such a chart; see attached (page 2). Don't look too closely at the ISPs; Merlin and SSME have relatively poor vacuum ISP because as first-stage engines they are optimized for sea-level ISP. The fuels aren't shown explicitly on the chart but the ISP indicates the fuel.

That and you have to be careful to separate engines that are actually fully developed and flying from "paper" engines. The SSME is flying. The RL-10 is flying. The J-2X isn't. Upgraded RS-68 engines are in development, but aren't flying.

One way that you could increase J-130's performance by a coupla mT - have specific 3-engine & 4-engine thrust structures & piping.

J-130 programme should not be affected in either cost or schedule, but there would be additional costs for phase 2 / J-24x development - substantial costs, I'd guess.

What you say is true, especially the cost part. Developing two separate thrust structures and piping not only increases development costs, but would also likely increase operational costs, because now you've got to manufacture two different thrust structures. To keep all of this straight will cost more money. You don't want a component to be accidentally manufactured to the wrong specs. This could potentially lead to two sets of tooling, procedures, and etc.

Again, I think Direct has it right. Develop one launch vehicle and just launch the core stage plus SRB's minus one main engine. The money you'll save in manufacturing and operations costs are a bonus on top of the development cost savings.

Can any of the team characterize the sort of information (if any) the Commission is requesting? I'd think you could work out what their line of thinking was on DIRECT by looking at the questions or requests they submit.

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More fundamental question, have they requested anything else? Is there a timeline for when they want additional information? I know the analysis teams are already cranking on their hamster wheels and abaci, so it would have to be soon to have an impact.

During our available time, we strongly made the point that, being an integrated architecture, the most significant driver that sizes much of the architecture is lunar global access. This is by far the most dominant driver in how much mass must be delivered to translunar injection. Indeed, with our present baseline, the size of the rocket and lander alone do not enable global lunar access - to attempt that would result in a rocket that is too large to reasonably build. That's just the physics of the problem. So we utilize all of the parameters at our disposal (lander propulsion load, loiter time in lunar orbit, etc) to open the hardest to reach places to exploration with what we considered a reasonable heavy lift launch vehicle to lift the necessary mass out of Earth's gravity well - and sure enough, several of the most interesting sites are in such locations. We also pointed out that we had scaled back our Ares V and Altair assumptions to supporting only equatorial and polar landings (while still protecting a reasonable level of cargo delivery capability for establishing an outpost) and the cost variance was only roughly 10 percent cheaper.

Not suggesting anything sinister here, and I've certainly seen this information presented from NASA before. I wonder what prompted to make this point so explicitly here?

Quote

Much attention has been focused on the probability of loss of crew (pLOC) as a figure of merit in determining the crew launch aspect of the architecture, and we expressed that the ESAS pLOC numbers were all using the same methodology and that the value was in the comparative results and not in the absolute numbers. Very simply, Ares' clear advantage is in the comparative simplicity of its first stage (the shuttle SRM) and use of a single gas generator cycle upper stage engine. These two attributes alone provide substantial robustness over, for example, a more complex liquid pump fed first stage and a multiengine upper stage - simply put, they are more complex with more moving parts. What Ares affords us, in accordance with the findings of the CAIB, is a crew launch system that has the potential to achieve unmatched safety in human spaceflight history. And this is not just a Constellation 'claim' as some would suggest, but has been validated by independent experts in the field of physics based probabilistic risk assessment. There will be much more provided on this topic as well.

(My highlighting).

Paraphrase - "DIRECT will never match Ares I's pLOC numbers".

I find it fascinating that 1x J2-X gives better numbers over a cluster of RL-10's with tolerance to a benign single failure.

Understand the point re solid vs multiple SSME's.

I watched a fascinating Apollo documentary last night, and interesting to see comments re the violence of takeoff on Apollo 8 - the guys actually thought they'd impacted the pad. How does that compare to expected levels of Ares I TO, if it happens?

Quote

On the topic of 'human rating', it is clear that the panel will want to hear more on this topic as well. The term gets thrown around in the community without a consistent understanding of what 'human rating' means. NASA's human rating 'policy' is clearly documented, but Constellation is the first program to really attempt to apply to a design in a practical manner. Our overall approach to human rating has been briefed to the ASAP, as has our program-wide approach to risk-based design that chooses robustness over blind fault tolerance in engineering these systems. All of our external review has largely validated this approach to date.

Paraphrase - "the mass changes as applied for the analyses are clearly documented in the HR policy, and have been briefed to ASAP".

Sorry to hear that NASA's analysis is once again flawed. Maybe the tools they're using can't model the upper stage properly. Anyway, I seem to recall a Popular Mechanics article on Direct where an aerospace executive (Bernard Kutter according to the article) went on record that ULA could indeed produce the upper stage. Has that fact been made known to the Commission? Perhaps ULA's findings and independent reviews confirming your findings would help counter NASA's claims.

I watched a fascinating Apollo documentary last night, and interesting to see comments re the violence of takeoff on Apollo 8 - the guys actually thought they'd impacted the pad. How does that compare to expected levels of Ares I TO, if it happens?

At no point in the history of Mercury, Gemini or Apollo did the crew ever experience greater vibration forces that +/- 0.6 g.

Even with the mitigation efforts, TO on Ares-I is expected to still be able to impart up to +/- 2.0g of vibrations on the Crew Module, although seat isolators are hoped to reduce that for the crew themselves.

Even a nominal Ares-I/Orion mission would still routinely subject the crew to the maximum vibration environment experienced on any previous system (0.6 to 0.8g).

Ross.

« Last Edit: 06/25/2009 07:28 PM by kraisee »

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Even with the mitigation efforts, TO on Ares-I is expected to still be able to impart up to +/- 2.0g of vibrations on the Crew Module, although seat isolators are hoped to reduce that for the crew themselves.Ross.

At those kind of alternating loads you have to start looking at Metal fatigue.. not just Ultimate and Tensile strength. I would want a higher FS for material in that envirnoment.

Since the Direct baseline architecture is 2 J-246's, with two JUS's, one fully fueled, one fueled just enough for LEO of CEV/LSAM.

Could both JUS's be partially fueled, and have a rendevous in LE1 or LLO? You have to expend two anyway, couldn't the docking procedures be simplified that way?

Also, in that vein, if a Centaur or D4US (are they the same? Don't know much about them, if they are two terms for the same upper stage or not) can take Orion around the moon for an Apollo 8 type flyby, could one be used to take the LSAM there too, again for a LE1 or LLO rendevous?I know fueled LSAM is heavier than Orion, so I don't know.But now I'm curious.

You've just stumbled upon what, in my view, is the biggest weakness of the EOR-LOR mission profile. Try as you might, you can't distribute the total payload evenly among multiple identical launch vehicles without on-orbit propellant transfer.

If the DIRECT ethos can be distilled to "one kind of vehicle launched multiple times", then EOR-LOR is a questionable mission profile, a holdover from the 1.5-launch approach that doesn't make as much sense in a 2-launch architecture.

The DIRECT 3.0 architecture calls for one launch of about 100 mT and another of about 70 mT. It's not really a 2-launch architecture, hence the heavily-offloaded J-24x CLV and the barely-viable J-130 CLV alternative.

With LOR-LOR, L1R-L1R, or (especially) L2R-L2R, the CEV and LSAM each have their own upper stage for TLI and insert themselves into rendezvous orbit separately.

The key thing to understand is that the LSAM actually masses less than the CEV when it separates for lunar descent. The LSAM is only heavier than the CEV at liftoff because it does the LOI burn for itself and the attached CEV.

Remember, the Apollo CSM was much more massive than the LM, mostly because the CSM did the LOI burn for the combined mass. Whichever spacecraft does LOI becomes much bigger than the other.

But if both spacecraft do LOI and their rendezvous masses are similar, then their TLI masses are similar, and therefore their LEO requirements are similar, and they can be lofted on identical launch vehicles.

With LOR-LOR, the rendezvous mass of the CEV is notably higher than the LSAM, because now it has to do its own LOI instead of relying on LSAM. So the CEV drives launch vehicle requirements.

But with L1R-L1R or (especially) L2R-L2R, the CEV takes a cheaper round trip to the rim of the moon's gravity well. This increases LSAM mass, but it decreases CEV mass by a much greater amount, and the combined effect helps even out the rendezvous masses.

With L2R-L2R, CEV liftoff mass is roughly the same as with EOR-LOR (depending on trajectory), even though it does its own LOI, and LSAM liftoff mass is dramatically reduced to about 20 mT, not much less than the CEV.

Either J-130 or Not Shuttle-C could lift a 25 mT spacecraft with a 45 mT EDS to put it through TLI. With a 2-launch L2R-L2R profile, this is enough for the baseline lunar mission.

Additionally, this same 5m Centaur-derived EDS could double as the new upper stage for EELV, and it would only make sense for ULA to lead the development, rather than NASA/MSFC.

This mission profile allows for the development of a significantly smaller upper stage that's much more versatile and would see higher flight rates. It also allows for the development of a significantly smaller LSAM descent stage with a lower center of gravity for landing stability.

The number of SSMEs expended per mission is reduced to six, global access to the lunar surface without expensive plane-change maneuvers, and global communications relay to earth via CEV at EML2.

If you try to replicate a 1.5-launch EOR-LOR architecture with one kind of launch vehicle, the closest you can get is DIRECT. But for a true 2-launch architecture, L2R-L2R makes more sense.

Butters,

Ok, you now make me ask a lot more questions than the first time. To help me understand what you just described better, maybe you can explain a few of your comments.

Briefly, what are the advantages/disadvantages to renevous at EML1, EML2, and lunar orbit. Apollo went to orbit, correct?I can understand EML1, as you are between the Earth and Moon. For L2R, you'd need to be on the far side of the moon? Wouldn't that burn extra fuel to get there than L1R?

In either L1R or L2R, the CSM would stay there, and the LSAM would decend to LLO, with the descent stage performing LOI, and then firing more to degrade that orbit slowly until you're over your landing zone, correct?And the advantage with this is the CSM doesn't need to escape lunar gravity, so it doesn't need that fuel? It just nudges itself out of the lagrange point and falls back to Earth? How would that work at the L2 point? It'd fall to the moon would it not? Would it do a swing by to the L1 point, then fall to Earth? Seems like it'd need some fuel for that, that the L1 points wouldn't require.

Why did Apollo take the CSM and LEM to LOI?How do you come to a stable position at a lagrange point? Does it require an insertion burn like orbit, or do you escape with just enough velocity so you coast to a stop right at the lagrange point and require no braking burn?

Just not sure how that works, so when you are whipping out terms about masses to rendevous, etc. I was a bit lost.

So when I understand that better, that will help be understand your ideas better.So, are you saying that if the CSM and LSAM were each launched separately, with an L2 rendevous, they could each utilize a 5m Centaur? And the LSAM mass could be reduced from 45mt to 20mt because it doesn't need to perform LOI for itself and the CSM? Would it still be as capable? Would lit only be 25mt lighter of propellent? That seems like a lot.

So what you are saying is that by launching the LSAM and CSM separately, and having them rendevous at the L2 point, you could use the already developed Centaur, Orion and Altair with as much hardware and as capable as the current Ares baseline, and launch them on a Pair of J-130's?

So why isn't that the baseline???Is the rendevous and docking at L1 or L2 complex and dangerous or something?Otherise, why worry about a new 8.4m upper stage? And the somewhat tricky docking in LEO of the LSAM and CSM with the 8.4m EDS?